1. Field of the Invention
The present invention relates generally to an air cooled turbine airfoil, and more specifically to a shaped film cooling hole in the airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with multiple stages of stator vanes and rotor blades that react with a hot gas flow to drive the engine and produce power. The turbine airfoils are exposed to such high temperatures that thermal damage would occur if not for the application of internal and external cooling air. The cooling of airfoils includes convection cooling, impingement cooling and film cooling in the airfoils exposed to the highest temperatures such as the first stage and even second stage airfoils.
Film cooling is produced by discharging the pressurized cooling air from the internal cooling passages onto the airfoil external surface. This creates a protective layer of film air to protect the metal airfoil surface from the hot gas flow. Prior art film holes include the straight circular entrance region having a constant diameter followed by a single conical diffusion section that opens onto the airfoil surface. The constant cross section entrance region is used for metering the cooling air flow through the film hole. The conical diffusion section is used for reducing the cooling air momentum or exit velocity of the air. If the air flow is discharged at too high of a velocity or at too high of an angle with respect to the airfoil surface, no film layer will develop.
Normally, an expansion area ratio of 2 to 6 times the metering section area is used in the airfoil film hole cooling design. This type of film cooling hole construction can be found in most of the prior art turbine airfoil cooling designs. FIG. 1 shows the prior art Vehr hole with a standard 10×10×10 shaped diffusion hole that is widely used in the current cooling designs for airfoils. See U.S. Pat. No. 4,653,983 issued to Vehr on Mar. 31, 1987 and entitled CROSS-FLOW FILM COOLING PASSAGES. The diffusion section has a 10 degree spanwise expansion in both the two side walls and the downstream expansion, while the upstream wall is straight and without an expansion. An expansion in the upstream direction will entrain hot gas into the film cooling hole at the exit plane as indicated by point A in FIG. 3. As a result, the entrainment causes shear mixing with the ejected cooling air and a degradation of the film effectiveness level. The 10×10×10 shaped diffusion holes are currently produced by the well-known EDM or electro-discharge machining process. FIG. 1 shows a cross sectional view of the hole, FIG. 2 shows a top view, and FIG. 3 shows a gun barrel view of the hole. FIG. 4 shows the EDM electrode that is used to produce the film cooling hole. As indicated from the top view of FIG. 2, the foot print is in the trapezoidal shape with four sidewalls. The same geometric shape is shown for the gun barrel view in FIG. 3. The metering hole circle is tangent to the upper or upstream side wall of the trapezoid.
As the TBC (thermal barrier coatings) technology improves, industrial gas turbine (IGT) airfoils can be applied with a thicker TBC. Machining film cooling holes using the EDM process becomes less cost effective. Since the TBC material is a non-conducting material (typically a ceramic), the electrode will not be able to cut through the TBC material to form the holes. Film cooling holes must be machined before the TBC can be applied. Thus, masking of the film cooling holes is required before the TBC can be applied. Then, the masking material is removed to leave the open holes in the TBC. This is a very costly and highly laborious process to form an airfoil with a TBC and film cooling holes.